1. Field of the Invention
The invention relates to a flight control apparatus for helicopters for use in driving a steering mechanism on the basis of the amount of control performed by the helicopter""s pilot.
2. Description of the Related Art
In flight of a helicopter, it is necessary to perform flight control with respect to four axes in total, i.e., a CP (collective pitch) axis corresponding to the thrust of a main rotor, a pitch axis corresponding to the inclination in the longitudinal direction of the airframe axis, a roll axis corresponding to the inclination in the transverse direction of the airframe axis, and a yaw axis corresponding to the heading (the thrust of a tail rotor). The pilot has to operate right/left hand control levers and right/left foot control pedals to control the direction of flight and the attitude of the helicopter while keeping the four axes in balance. Accordingly, a very sophisticated control technique is required of the pilot.
FIG. 8 is a diagram showing the configuration of an example of the prior art flight control apparatus for helicopters. This apparatus is a mechanical control transmission mechanism adopted in most of the existing helicopters. A control unit 1 including control levers and control pedals is coupled to a steering mechanism 4 for controlling the pitch angle of a blade, via a mechanical link mechanism 2, in which at some midpoint an actuator 3 for increasing the control force is interposed.
In such a mechanical control transmission mechanism, the amount of control in the control unit 1 is directly affected by the amount of driving of the steering mechanism 4. Accordingly, in some cases even a slight control error may largely disturb the balance among the four axes, so that a relatively heavy burden is put on the pilot. As a countermeasure, a stability augmentation system (SAS), an automatic flight control system (AFCS) or the like is added to the link mechanism 2, so that the flight stability and control response property of the helicopter are improved.
FIG. 9 is a diagram showing the configuration of another example of the prior art flight control apparatus for helicopters, which has an electrical control transmission mechanism under development. In the electrical control transmission mechanism, which is generally referred to as a fly by wire (FBW) system, a control unit 1 including control levers and control pedals is coupled to a steering mechanism 4 for controlling the pitch angle of blade, via wires for transmitting electric signals. At some midpoint are interposed an amount-of-control sensor 5 for detecting the amount of control, a computer 6 which carries out operation of various sensor signals to output a driving signal for each axis, and an actuator 3 for driving the steering mechanism 4 based on the driving signals.
In such an electrical control transmission mechanism, referring to the amount of control performed by the pilot, the computer 6 can generate driving signals optimum for the helicopter. In addition, in case where an unstable control signal is inputted due to a failure of a component, the computer 6 can remove the signal.
The mechanical control transmission mechanism has a relatively simple construction as compared with the electrical type and, accordingly, has a lower system failure rate than the electrical type (for a case where the failure rates are compared in the systems with the same redundancy (multiplicity)). As a result, the reliability of the mechanical type is higher. Because of the advantages, many mechanical types are practically used. However, the response type of the helicopter to the amount of control performed by the pilot is restricted to one type (i.e., a type in which the changing rate of airframe attitude is controlled). Therefore, the manner of stabilization by means of the stability augmentation system or automatic flight control system is limited to providing the function of adding dumping to the rate or of simply maintaining a constant attitude.
On the other hand, the electrical control transmission mechanism can realize any input and output system and calculation model using a computer. Thus, the response type of the helicopter to the amount of control performed by the pilot can be selected from among a plurality of flight control laws, so that the diversity of control modes can be increased. In addition, by performing optimum control in each control mode, the flight stability and control response property of the helicopter can be dramatically improved. However, the electrical control transmission mechanism is essentially different from the mechanical type, and therefore, it is difficult to practically use the electrical type in existing helicopters by modification, although it is possible to adopt the electrical type to newly developing helicopters.
It is hence an object of the invention to provide a flight control apparatus for helicopters, which can be easily applied to mechanical control transmission mechanisms that are mounted on many existing helicopters, and which can take advantages of the above-described characteristics of the electrical type control transmission mechanism, thereby remarkably increasing the characteristics and performance of the existing helicopters.
The invention provides a flight control apparatus for helicopters, comprising: a control unit operated by a pilot; a steering mechanism for generating an aerodynamic control force; a control transmission mechanism for mechanically transmitting an amount of control Ma in the control unit to the steering mechanism, to thereby drive the steering mechanism; an amount-of-control sensor for detecting the amount of control Ma in the control unit to output a control signal Sa; a flight control law calculation unit for calculating a flight control law of a helicopter based on the control signal Sa to output a driving signal Sb for the steering mechanism; a difference calculation unit for subtracting the control signal Sa from the driving signal Sb to output a difference signal Sc; and a servo actuator unit for adding an amount of difference Mc corresponding to the difference signal Sc, to the amount of control Ma transmitted via the control transmission mechanism.
According to the invention, the amount of control Ma in the control unit is transmitted to the steering mechanism via the control transmission mechanism, so that the same operation as that of the prior art mechanical control transmission mechanism is performed. In addition, the amount-of-control sensor detects the amount of control Ma in the control unit, and supplies the control signal Sa to the flight control law calculation unit for calculating the flight control law of the helicopter.
In the same manner as the prior art electrical control transmission mechanism, the flight control law calculation unit can perform a model calculation corresponding to various flight control laws by using the full operational range (full-authority) of the control system based on the control signal Sa. A result of the calculation is outputted as a driving signal Sb for the steering mechanism, and the driving signal Sb is a driving signal optimum for the helicopter. However, in the case where the driving signal Sb is directly supplied to the servo actuator unit, such a problem occurs that the amount of control Ma, which has been already supplied to the steering mechanism via the control transmission mechanism, serves as an excess with respect to the amount of driving Mb of the steering mechanism. Thus, the difference calculation unit subtracts the control signal Sa from the driving signal Sb, and supplies the obtained difference signal Sc to the servo actuator unit, so that the amount of driving Mb of the steering mechanism does not include the excess amount of control Ma.
Accordingly, part of the amount of driving Mb of the steering mechanism reflects the amount of control Ma in the control unit, and the remaining part of the amount of driving Mb reflects the calculation result of the flight control law calculation unit. According to the invention, while the reliability of the prior art mechanical control transmission mechanism is always maintained, sophisticated technological flight control laws by the prior art electrical control transmission mechanism can be applied. The characteristics and performance of the control transmission mechanism are remarkably enhanced as a whole, and the flight stability and control response property of the helicopter are dramatically improved.
When relationship between a mechanical amount M and an electrical amount S is represented by a function of M=f(S), the relationship in the control unit is represented by a function of Ma=f(Sa), and that in the servo actuator unit is represented by a function of Mc=f(Sc). Since the relationship of Sc=Sbxe2x88x92Sa is established, the amount of driving Mb of the steering mechanism is obtained as follows:                     Mb        =                  xe2x80x83                ⁢                  Ma          +                      M            ⁢                          xe2x80x83                        ⁢            c                                                  =                  xe2x80x83                ⁢                  Ma          +                      f            ⁡                          (              Sc              )                                                              =                  xe2x80x83                ⁢                  Ma          +                      f            ⁡                          (                              Sb                -                Sa                            )                                                              =                  xe2x80x83                ⁢                  Ma          +                      f            ⁡                          (              Sb              )                                -                      f            ⁡                          (              Sa              )                                                              =                  xe2x80x83                ⁢                  f          ⁡                      (            Sb            )                              
From the above, it will be seen that the amount of driving Mb of the steering mechanism coincides with the driving signal Sb of the flight control law calculation unit.
Although the description concerns the control for one axis, the invention can be applied to the control relating to all or part of four axes, i.e., a pitch axis, a roll axis, a yaw axis, and a CP axis.
In a mechanical control transmission mechanism which is mounted on many existing helicopters, a flight control function equivalent to the FBW system can be easily realized only by performing modification of adding the amount-of-control sensor, flight control law calculation unit, difference calculation unit and servo actuator unit according to the invention.
In the invention it is preferable that the servo actuator unit is disposed functionally in series to the control transmission mechanism.
According to the invention, the process of summing the amount of control Ma transmitted via the control transmission mechanism and the amount of difference Mc outputted from the servo actuator unit can be easily realized.
In the invention it is preferable that in the event a failure of the flight control law calculation unit occurs, the servo actuator unit is locked after eliminating a deviation Mc between the amount of control Ma and the amount of driving Mb of the steering mechanism.
According to the invention, when the flight control law calculation unit is normal, the sum of the amount of control Ma from the control unit and the amount of difference Mc from the servo actuator unit is supplied to the steering mechanism, so as to properly maintain the amount of driving Mb. When the flight control law calculation unit fails, there arises a deviation between the amount of control Ma and the amount of driving Mb, corresponding to the amount of difference Mc immediately after the failure. Therefore, the pilot is impelled to continuously perform the control in a condition where the pilot feels the deviation between the position sense of the control portion (levers and pedals) in the control unit and the controllable range of the helicopter. This exerts a heavy burden on the pilot. To solve this problem, in the event a failure of the flight control law calculation unit occurs, the operating section of the servo actuator is once locked under the condition that the deviation Mc between the amount of control Ma and the amount of driving Mb of the steering mechanism is kept fixed. The pilot is allowed to perform the initial counter measure control under this state, so as to assure safety. The lock is released by an instruction (switching) by the pilot after the correction control to eliminate the deviation Mc. Thereafter, the operating section of the servo actuator unit is again locked. The position sense of the control portion in the control unit thus agrees with the controllable range of the helicopter. The pilot can continue the control by the same control method as that in the prior art mechanical control transmission mechanism, so that the burden on the pilot can be largely reduced.
In the invention it is preferable that a power boost unit having an SAS function which effectively functions only in the event a failure of the flight control law calculation unit occurs is disposed between the servo actuator unit and the steering mechanism.
According to the invention, even in the case where the use of the flight control law calculation unit is stopped because of a failure or the like, the control transmission mechanism SAS (stability augmentation system) function can be incorporated therein as a substitute function, so that the flight stability and control response property of the helicopter can be continuously ensured.